Integrated control for a turbopropulsion system

ABSTRACT

An integrated turbopropulsion system is disclosed consisting of a variable air induction system, a turbine engine and an exhaust nozzle/ejector. Control of the total system is provided by a plurality of basic control systems consisting of separate controls for each individual system variable. A supervisory control receives inputs from selected basic controls and from external signals, and provides additional inputs to the basic controls to permit the basic controls to increase the performance of the turbopropulsion system and to extend the control envelope of the total system. Additional basic controls may be added to the system, or basic controls may be subtracted from the system without affecting the performance of the entire turbopropulsion system. Novel basic controls for individual system variables are disclosed, as well as a novel integrator lockout control and a novel variable rate control. As a result of the intercommunication between individual basic controls provided by the supervisory control, the entire integrated turbopropulsion system may be controlled to more stringent limits than can be provided by a plurality of independent basic controls.

United States Patent 1191 Webb et al. Mar. 19, 1974 [54] INTEGRATED CONTROL FOR A 3,482,396 12/1969 Nelson et al. 0 33 3 TURBOPROPULSION SYSTEM 3.469.395 9/1969 Spitsbergen et a1 60/3928 R Primary Examiner-Carlton R. Croyle Assistant Examiner-Robert E. Garrett Attorney, Agent, or FirmDonald F. Bradley all of Fla-1 [57 ABSTRACT [73] Assignee: United Aircraft Corporation, East An integrated turbopropulsion system is disclosed Hartford, Conn. consisting of a variable air induction system, a turbine engine and an exhaust nozzle/ejector. Control of the [22] Filed June 1973 total system is provided by a plurality of basic control [21] Appl. No.: 374,611 systems consisting of separate controls for each indi- Relate'd US. Application Data yidual system variable. supervisory control receives inputs from selected basic controls and from external [63] Continuation of Ser. No. 138,163, Apr1l28, 1971. Signals, and provides additional inputs to the basic controls to permit the basic controls to increase the [52] 60/39'16 60/3928 60/226 performance of the turbopropulsion system and to ex- 60/236* 60/238 60/239 60/240 tend the control envelope of the total system. Addi- [511' ll'it. Cl F02C 9/04, FOZC 3/06 tional basic controls may be added to the System or [58] Flew of Search 60/3928 3928 39-16 R basic controls may be subtracted from the system without affecting the performance of the entire turbo- [561 References cued propulsion system. Novel basic controls for individual UNITED STATES PAT NTS system variables are disclosed, as well as a novel inte- 3.s20.133 7/1970 Loft et al. 60/3928 R x g r l cko ontrol n a novel variable rate con- 3.606.754 9/1971 White 60/3928 R trol. As a result of the intercommunication between 3.523.423 8/1970 Young 60/39- R X individual basic controls provided by the supervisory 3.283.499 11/1966 Scheidler R control the entire integrated turbopropulsion ystem {aplm 55'553 may be controlled to more stringent limits than can be 8181' et a 3.307.353 3/1967 Stearns 60/3928 R provlded by a plurahty of Independent ham controls 3.639.076 2/1972 Rowen 60/3928 T X 15 Claims, 20 Drawing Figures /fl /f i. zzifi j 2.2 f

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sum :15 or 16 INTEGRATED CONTROL FOR A TURBOPROPULSION SYSTEM This is a continuation of application Ser. No. 138,163, filed Apr. 28, 1971.

The invention herein described was made in the course of or under a contract with the Department of the Air- Force.

BACKGROUND OF THE INVENTION 1. Field of the Invention This invention relates to an integrated turbopropulsion system in which the control of a variable air induction system, an exhaust nozzle/ejector and a turbine engine are integrated. Individual basic controls are provided for each of the individual system variables. In addition, a supervisory control provides intercommunication between selected basic controls to permit the basic controls to operate closer to the limits of the turbopropulsion system and insure optimum system operation.

2. Description of the Prior Art Historically the individual propulsion system components, namely, inlet, engine and ejector, have been procured separately and assembled by the aircraft manufacturer. In addition, controls for each of the individual components have also been procured separately. Intercommunication between the propulsion system controls was kept to a minimum to reduce the interaction between components, and to maintain the responsibility of the manufacturer for each component.

Advanced propulsion systems for present and future aircraft range from subsonic to supersonic flight speeds. For maximum system performance and effectiveness, the entire propulsion system must operate efficiently over the entire flight envelope of the aircraft, and continuous operation at or near physical operating limits requires improved propulsion system performance under severe environmental conditions.

The present integrated turbopropulsion system provides improved performance in that the propulsion system components are totally integrated in a single package, and the controls for each of the system components are fully integrated by a supervisory control with intercommunication being freely transmitted between the basic control systems to insure optimum system operation.

The turbopropulsion system described herein consists of an inlet, engine and ejector. The inlet is an ex-' ternal internal compression air induction system with variable throat geometry, variable spillage bypass doors, and variable secondary airflow doors. The engine is a twin spool, split flow, duct heating turbofan with variable compressor bleed, variable compressor guide vanes, variable duct nozzle, variable low turbine bypass, and variable gas generator nozzle area. Two fueled regions are the gas generator combustor and the duct heater. The ejector consists of two variable areas and a mixing volume for the flow streams. The two variable areas are blow-in doors and a variable area ejector nozzle. Control of the entire turbopropulsion system provides improved performance by matching inlet and engine airflow, improving fuel consumption, and providing additional flow and control stability margin as well as additional performance and structural margin during critical operations.

A basic control system is provided for each individual system variable. In addition, a supervisory control provides integration of the entire turbopropulsion system control. The supervisory control receives input signals from various of the individual basic controls, and also receives external signals concerning aircraft operating conditions. From these various inputs, the supervisory control produces additional data for the basic control systems, adding to or subtracting from the basic controls in a manner to increass the system control envelope for the entire turbopropulsion system over and beyond the envelope provided by the basic control. In other words, the basic control, with additional inputs from the supervisory control, will operate the turbopropulsion system variable under its control closer to the limits which provide maximum performance than can the basic control itself. As a result, both performance and stability margin of the entire system are improved.

One advantage of the present invention involving the control of a turbopropulsion system by a supervisory control is that basic controls may be added to or subtracted from the system without radically affecting the overall control. The job of the supervisory control is to coordinate the control provided by each of the basic controls to the entire turbopropulsion system for the specific purpose of providing maximum system performance.

Different degrees of authority for an individual system variable may be assigned to the basic and/or supervisory control. It may be desired to control a variable only by the basic control, without any input from the supervisory control. On the other hand, if complete control of a variable is accomplished only by the supervisory control, this variable will be maintained at only one position if the supervisory control is removed or is rendered inoperative. In the present system, the supervisory control may be inactivated or removed, and the basic control systems provide a level of performance for each of the system variables to produce adequate, though less than optimum, performance of the turbopropulsion system without the supervisory control.

In addition to the supervisory control system disclosed herein, there is also disclosed novel basic controls for selected variables in a turbopropulsion system.

Additional novel features of the present invention include an integrator lockout control which insures that the control system does not extend its operation beyond the limits of the system controls, and a variable rate control which changes the compensation of a system variable depending on the direction and distance from alimit of the variable.

SUMMARY OF THE INVENTION In accordance with the present invention, there is provided an integrated turbopropulsion system consisting of a variable air induction system, a turbine engine and an exhaust nozzle/ejector. A basic control system is provided for each variable required to be controlled. In addition, there is provided a supervisory control which receives inputs from the basic controls and also receives external signals, the supervisory control then providing additional inputs to the basic controls to optimize the operation of the basic controls and to insure optimum control of the entire turbopropulsion system. Each basic control is sufficient to provide acceptable, though less than optimum, operation if the supervisory control is inactivated. Basic controls may be removed from the system, or additional basic controls added, depending on the precise configuration of the turbopropulsion system, without necessitating comprehensive changes in the supervisory control system.

In accordance with other aspects of the present invention, there are provided novel basic controls for individual system variables. These include a novel fan surge margin control, a novel gas generator fuel control system, a novel turbine inlet temperature control system, a novel gas generator rotor speed limit control, a novel surge margin limit control, and a novel low turbine bypass area control.

Other embodiments of the present invention include a novel integrator lockout control which utilizes logic to insure that a control system does not extend its operation beyond the limits of the system controlled.

A further embodiment of the present invention incorporates a novel variable rate control which adjusts the gain of a compensation network as a function of the direction of movement and distance from a limit of a controlled variable.

BRIEF DESCRIPTION OF TI-IE DRAWINGS FIG. 1 is a schematic diagram of the integrated turbopropulsion system including the control variables.

FIG. 2 shows in block diagram form the basic controls for the variables of the turbopropulsion system of FIG. 1, together with the interaction of the supervisory control with the basic controls.

FIG. 3 is a block diagram showing the details of the inlet geometry control of FIG. 2.

FIG. 4 is a block diagram showing the compressor geometry control of FIG. 2.

FIG. 5 is a block diagram showing additional details of the gas generator fuel control of FIG. 2.

FIGS. 6 through 13 show in partially schematic and partially block diagram form the specific details of the gas generator fuel control of FIG. 5.

FIG. 14 shows in block diagram form additional details of the duct heater fuel control of FIG. 2.

FIG. 15 shows in block diagram form additional details of the low turbine bypass control, duct heater nozzle control and gas generator nozzle control of FIG. 2.

FIGS. 16 through 18 show partially in schematic form and partially in block diagram form specific details of the engine airflow control of FIG. 15.

FIG. 19 shows schematically the details of the integrator lockout control.

FIG. 20 shows schematically the variable rate control shown in block diagram form in FIG. 12.

DESCRIPTION OF THE PREFERRED EMBODIMENT The Turbopropulsion System FIG. 1 shows the integrated turbopropulsion system which is to be controlled. The turbopropulsion system consists of an inlet, an engine and an ejector.

The inlet 10 is an external-intemal compression air induction system with a variable throat geometry controllable by movement of the inlet spike 12. The inlet 10 also contains variable spillage bypass doors 14, and variable secondary airflow doors 16.

The engine 18 is a twin spool, split flow, duct-heating turbofan with variable compressor bleed 20, variable compressor guide vanes 22, a variable duct nozzle 24, a variable low turbine bypass 26 and a variable gas generator nozzle area 28. Two fueled regions in the turbopropulsion system are the gas generator combustor 36 and the duct heater 38.

The ejector 30 consists of two variable areas, blow-in doors 32 and a variable area ejector nozzle 34, and a mixing volume.

The components of the turbopropulsion system including the inlet 10, engine 18 and ejector 30 are well known in the art. Likewise, the controlled variables and the means for controlling each variable are well known to those skilled in the art. It is also apparent that the teachings of the present invention may be applied to other turbopropulsion systems which contain additional variable elements, or from which variable elements have been eliminated depending on the application of the particular turbopropulsion system.

The Supervisory and Basic Control Systems FIG. 2 shows the relationship of the basic control systems for each of the turbopropulsion system variables, and their relationship of the supervisory control with respect to the basic control systems. The basic control systems are shown within a block labeled 40 and comprise a plurality of individual basic controls. The supervisory control 42 communicates with the basic control 40 as will be described in detail, receiving information from the basic control 40 and providing additional inputs to the basic control 40. Supervisory control 42 also receives external signals shown at 44. Not shown in FIG. 2 are the various inputs such as temperature, pressure, etc. which are fed from the turbopropulsion system to each of the individual controls which comprise basic control system 40 and on which each of the individual basic controls operates.

The basic control system 40 comprises the following individual basic controls. Inlet geometry control 46 operates to provide maximum inlet pressure to the face of engine 18 while minimizing pressure distortions and perturbations. Control of the inlet spike position 12 (X,), inlet bypass door position 14 (Asp) and secondary airflow door position 16 (A are controlled by the inlet geometry control 46. Inputs to the inlet geometry control 46 include throat pressure ratio, actual spike position, shock pressure ratio, and aircraft attitude, acceleration and Mach number. The inlet geometry control will be described in detail in connection with FIG. 3.

The compressor geometry control 48 includes control of the compressor bleed 20 (CH) and control of the position of the variable compressor vanes 22 (CVA). The compressor bleed and compressor inlet guide vanes are both scheduled as a function of compressor speed and compressor inlet temperature. The compressor geometry control is described in additional detail in connection with FIG. 4.

-The gas generator fuel control 50 generates an error signal (W;) which controls the fuel flow by controlling the position of the fuel metering valve. This is the most complex of the basic control systems, and is described in further detail in connection with FIGS. 5 through 13. In general, seven error signals are generated and compared, with the error calling for the least fuel flow change selected to control the metering valve.

The duct heater fuel control 52 has the primary function of controlling the flow of fuel into the duct heater 38( Inputs to the duct heater fuel control include the duct inlet temperature, duct nozzle area, and power lever angle. Additional functions of the duct heater control are control of power lever rate and request of nominal duct nozzle area. The duct heater fuel control 52 is described in additional detail in FIG. 14.

The low turbine bypass control 54 controls the low turbine bypass area 26 (LTBA). This control is utilized to control gas generator speed, and thus control gas generator airflow. The duct heater nozzle control 56 controls the duct nozzle area 24 (DNA), which is used to control fan surge margin by providing a trim on fan airflow. The gas generator nozzle control 58 controls the gas generator nozzle area 28 (GGNA), which is utilized to control low rotor speed, thus providing control of total engine airflow. The low turbine bypass control 54, the duct nozzle control 56 and the gas generator nozzle control 58 are incorporated in the engine airflow control system described in additional detail in FIGS. l5, 16, 17 and 18.

The Supervisory Control System The basic functions of the supervisory control 42 with respect to the basic control systems incorporated in block are also shown in connection with FIG. 2. l

The details of.the interrelationships between the various basic controls and the supervisory control 42 will be explained in detail in connection with subsequent drawings.

The inlet geometry control 46 supplies an inlet bypass door area position signal (A through line 60 to the supervisory control 42. This signal passes through the supervisory control as shown by the dotted lines and is transmitted via line 61 to the gas generator nozzle control 58. This signal on line 61 represents the desired engine airflow biased by the actual inlet airflow.

The duct heater fuel control 52 supplies an inlet geometry reset signal (A Reset) through line 62 to the supervisory control 42 where it is combined with the external signals 44 fed into the supervisory control and transmitted to the inlet geometry control 46 through line 63. The external signals 44 supplied to the supervisory control are Mach number (M,,), pitch (a) and yaw (B). A rate limited power lever angle signal is fed from duct heater fuel control 52 to the gas generator fuel control via line 73.

The gas generator fuel control 50 transmits a low rotor (N speed limit signal through line 64 to the gas generator nozzle control 58 where it is continuously compared to a, requested low rotor speed (N to insure safe operating speed. Gas generator fuel control 50 also transmits a high rotor (N speed limit signal through line 65 to the supervisory control 42 where it is compared to the desired high rotor speed and transmittedvia line 66 as a resultant low turbine bypass area trim signal to the low turbine bypass control 54. Supervisory control 42 transmits a compressor surge margin reset signal to the gas generator fuel control 50 via line 67 which limits fuel flowto maintain a desired surge margin plus reset. Supervisory control 42 also transmits a fuel flow trim signal to the gas generator fuel control 50 via line 68, this signal modulating fuel flow to maintain the desired turbine inlet temperature.

The duct heater fuel control 52 transmits an open loop duct heater nozzle area request signal (DNA*) to the duct heater nozzle control 56 via line 69. This signal, in the absence of trim signals, is the nozzle area request.

The low turbine bypass control 54 transmits a delta area signal (A LTBA) to the gas generator nozzle control 58 via line 70. This signal trims gas generator nozzle area minimizing interaction effects.

The supervisory control 42 also transmits a fan surge margin reset signal and a nozzle area trim signal to the duct heater nozzle control 56. The reset signal resets the area to increase fan surge margin, while the trim signal modulates the area to maintain a desired fan surge margin plus reset. These signals are transmitted via lines 71 and 72 respectively.

Additional functions performed by the supervisory control 42 will be described in connection with the description of the various basic control systems.

The Inlet Geometry Control The inlet geometry control is described in detail in connection with FIG. 3. The features of the inlet geometry control are: first, optimum inlet spike control by direct sensing of aircraft Mach number and throat pressure ratio; second, throat Mach number control by selective'positioning of inlet geometry; third, minimized shock disgorgement with inlet bypass area adjustment; and fourth, increased stability margin control by sensing aircraft acceleration and flight attitude.

The spike position (X,,) basic control positions the spike as a function of Mach number and maintains desired throat pressure ratio. Furthermore, the spike position basic control provides a Mach number switch that changes the control mode, and provides automatic restarting. The supervisory control provides inputs of Mach number and aircraft attitude (pitch and yaw) which are combined with signals indicative of spike position, aircraft acceleration and inlet unstart to a pressure ratio computer which produces a signal indicative of throat pressure ratio request. The throat pressure ratio request signal is compared with the throat pressure ratio signal in pressure ratio comparator 82 to produce a throat pressure ratio error which actuates spike position actuator 84 to control the spike position in the inlet.

Spike position control is accomplished by an open loop schedule at low Mach number, and a closed loop schedule at high Mach number. At high Mach numbers, position is modulated to maintain a throat pressure ratio (P,/P,).

The spike position control provides control of Mach number at the throat of the inlet to the desired value. The desired pressure ratio is biased or reset whenever additional margin is required. The biases are yaw, pitch and local acceleration/deceleration (Mach number 

1. A fuel control for a twin spool turbine engine having a high turbine for driving a high compressor and a low turbine for driving a low compressor, and including a burner for providing power for said high and low turbines comprising means responsive to selected parameters in said engine for generating a plurality of signals each of which is indicative of a fuel flow limit for said engine, at least limit selector receiving each of said fuel flow limits signals and passing therethrough the fuel flow limit signal calling for the lowest fuel flow, means for generating a fuel flow error signal indicative of the difference between a desired fuel flow and actual fuel flow, a minimum error selector receiving the fuel flow limit signal passed by said least limit selector and receiving said fuel flow error signal and passing therethrough the signal calling for the lower fuel flow, means for generating a signal indicative of a minimum fuel flow limit for said engine, a maximum error selector receiving said lower fuel flow signal and said minimum fuel flow limit signal and passing therethrough the signal calling for the maximum fuel flow, a variable gain circuit for varying the gain of the maximum fuel flow signal passed by said maximum error selector, means responsive to the fuel flow limit signal passed by said least limit selector and to the minimum fuel flow limit signal for scheduling the gain of said variable gain circuit, and means responsive to said maximum fuel flow signal as compensated by said variable gain circuit for controlling the flow of fuel to said combustor.
 2. A fuel control as in claim 1 in which said means responsive to selected parameters in said engine for generating a plurality of signals indicative of a fuel flow limit for said engine comprises means for generating a signal indicative of engine high compressor speed limit, means for generating a signal indicative of low compressor speed limit, means for generating a signal indicative of high turbine inlet temperature limit, means for generating a signal indicative of burner pressure limit, and means for generating a signal indicative of compressor surge limit.
 3. A fuel control as in claim 1 in which said means for scheduling the gain of said variable gain circuit includes means responsive to said fuel flow limit signal for scheduling a first gain factor, means responsive to said minimum fuel flow limit signal for scheduling a second gain factor, means responsive to the magnitude and polarity of said maximum fuel flow signal for selecting one of said first or second gain factors, and means for scheduling the gain of said variable gain circuit in response to said selected gain factor.
 4. A fuel control as in claim 1 in which said means for generating a fuel flow error signal comprises a variable power lever connected with said turbine engine, means for generating a first signal in response to the position of saId power lever, means for generating a second signal in response to the temperature at the engine compressor inlet, means responsive to said first and second signals for generating a scheduled fuel flow signal, means responsive to said first and second signals for generating a scheduled high compressor speed signal, means for measuring high compressor speed and generating a measured high compressor speed signal, means comparing said scheduled high compressor speed signal with said measured high compressor speed signal to generate a high compressor speed error signal, means responsive to said measured high compressor speed signal for generating a governor slope signal, means multiplying said high compressor speed error signal by said governor slope signal to produce a governor slope bias signal, means for adding said governor slope bias signal to said scheduled fuel flow signal to produce a fuel flow request signal, means for measuring actual fuel flow and generating a signal in response thereto, and means for comparing said measured actual fuel flow signal with said fuel flow request signal to generate said fuel flow error signal.
 5. A fuel control as in claim 1 in which said means for generating a minimum fuel flow limit signal comprises means responsive to burner pressure and to compressor discharge temperature for generating a desired minimum fuel flow signal, means for measuring actual fuel flow and generating a signal in response thereto, and means comparing said measured fuel flow signal with said desired minimum fuel flow signal to generate said minimum fuel flow limit signal.
 6. A fuel control as in claim 1 and including means responsive to the temperature rise across said burner for generating a signal indicative of the occurrence of a flameout in said engine. and means for disconnecting said maximum fuel flow signal and scheduling a fixed fuel flow signal to said engine when a flameout signal is present.
 7. A fuel control as in claim 2 in which said means for generating a compressor surge limit signal comprises means responsive to the temperature at the inlet to said compressor and to the speed of said high compressor for generating a signal indicative of desired compressor discharge airflow, means for measuring compressor discharge airflow and generating a signal indicative thereof, and means for comparing said desired and measured signals to generate a compressor surge limit signal.
 8. A control as in claim 7 and including means for generating a reset signal, and means for varying said desired compressor discharge airflow signal in response to said reset signal to provide additional surge margin under selected operating conditions of said engine.
 9. A fuel control as in claim 2 in which said means for generating a high compressor speed limit signal comprises means responsive to the temperature at the discharge of said high compressor and to the temperature at the inlet to said high turbine for generating a signal indicative of a desired high compressor speed limit, means for measuring high compressor speed and generating a signal indicative thereof, and means for comparing said desired high compressor speed limit signal with said measured high compressor speed signal to generate a high compressor speed limit signal.
 10. A control as in claim 9 in which said engine includes a variable bypass around the turbine, a control for said bypass, and means for feeding said high compressor speed limit signal to said bypass control to vary the bypass area.
 11. A control as in claim 9 in which said turbine engine includes a source of cooling air for said turbine, and means responsive to the cooling airflow for varying said desired high compressor speed limit signal.
 12. A fuel control as in claim 1 in which said engine has a source of cooling air for the turbines, and in which said means for generating a high turbine inlet temperature limit sigNal comprises means responsive to the rate of turbine cooling airflow and the temperature of the turbine cooling airflow for generating a desired high turbine inlet temperature limit signal, means for measuring the high turbine inlet temperature and generating a signal in response thereto, and means for comparing said measured high turbine inlet temperature signal with said high turbine inlet temperature limit signal to produce said high turbine inlet temperature limit signal.
 13. A control as in claim 12 in which said engine has a variable power lever connected therewith for scheduling engine operation, and means responsive to the position of said power lever for varying said desired high turbine inlet temperature limit signal.
 14. A fuel control as in claim 1 in which said means for generating a burner pressure limit signal comprises means responsive to the temperature at the discharge from said high compressor for generating an allowable burner pressure signal, means for measuring the pressure in said burner and generating an actual burner pressure signal, and means comparing said actual burner pressure signal with said allowable burner pressure signal to produce said burner pressure limit signal.
 15. A control as in claim 14 and including a duct surrounding the burner, means for measuring the pressure in said duct and generating a duct pressure signal, and means for varying said burner pressure limit signal in response to said duct pressure signal. 